Variable guide vane sealing

ABSTRACT

A variable guide vane assembly is provided for a turbine defining a core air flowpath. The variable guide vane assembly includes an airfoil band defining a flowpath surface and a cavity. The variable guide vane assembly further includes an airfoil including a first end extending at least partially into the cavity of the airfoil band and an opposite second end, the airfoil extending generally along an axis between the first end and the second end and being moveable generally about the axis relative to the airfoil band. The variable guide vane assembly further includes a sealing element operable to form a seal between the first end of the airfoil and the airfoil band.

CROSS-REFERENCE TO RELATED APPLICATIONS

This application is a continuation application of U.S. application Ser.No. 15/796,942 filed Oct. 30, 2017, which is hereby incorporated byreference in its entirety.

FEDERALLY SPONSORED RESEARCH

This invention was made with government support under contact numberFA8650-15-D-2501 awarded by the Department of the Air Force. The U.S.government may have certain rights in the invention.

FIELD

The present subject matter relates generally to gas turbine engines.More particularly, the present subject matter relates to sealingassemblies for variable vanes in gas turbine engines.

BACKGROUND

Gas turbine engines generally include a compressor section, a combustionsection, and a turbine section in serial flow order. The compressorsection may include one or more compressors, each of the one or morecompressors typically including sequential stages of compressor rotorblades and compressor stator vanes. Similarly, the turbine section mayinclude one or more turbines, each of the one or more turbines typicallyincluding sequential stages of turbine rotor blades and turbine statorvanes.

The stages of stator vanes in the one or more compressors and/or the oneor more turbines may change a direction of an airflow thereacross inorder to increase a performance and efficiency of the gas turbineengine. The performance and efficiency of the gas turbine engine mayfurther be increased by including stator vanes in the one or morecompressors and/or the one or more turbines capable of rotating about anaxis in order to vary a direction in which the stator vanes change theairflow thereacross. These are commonly referred to as a variable statorvanes.

Despite the increases in performance and efficiency derived from theinclusion of variable stator vanes in the one or more compressors and/orthe one or more turbines, in at least certain engines, at least aportion of the airflow thereacross may be capable of leaking around aradially inner end and/or a radially outer end of the variable statorvanes by virtue of the variable stator vanes not being fixedly attachedto a respective radially inner or radially outer band. Such may have adetrimental effect on the gas turbine engine's performance, efficiency,and durability.

Accordingly, a stator vane assembly capable of varying a direction inwhich it directs airflow thereacross while minimizing an amount ofleakage around its radially inner end/or radially outer ends would beuseful.

BRIEF DESCRIPTION

Aspects and advantages of the invention will be set forth in part in thefollowing description, or may be obvious from the description, or may belearned through practice of the invention.

In one exemplary aspect of the present disclosure, a variable guide vaneassembly for a machine defining a core air flowpath is provided. Thevariable guide vane assembly includes an airfoil band defining aflowpath surface and a cavity. The variable guide vane assembly furtherincludes an airfoil including a first end extending at least partiallyinto the cavity of the airfoil band and an opposite second end, theairfoil extending generally along an axis between the first end and thesecond end and being moveable generally about the axis relative to theairfoil band. The variable guide vane assembly further includes asealing element operable to form a seal between the first end of theairfoil and the airfoil band.

In certain exemplary embodiments the sealing element is positioned atleast partially within the cavity of the airfoil band.

For example, in certain exemplary embodiments the airfoil defines aleading edge and a trailing edge, wherein the first end of the airfoildefines an airfoil slot extending in a direction generally from theleading edge to the trailing edge, and wherein the sealing elementextends at least partially into the airfoil slot of the airfoil.

For example, in certain exemplary embodiments the airfoil defines apressure side and a suction side, wherein the airfoil slot of theairfoil extends from the pressure side to the suction side, and whereinthe sealing element extends through the airfoil slot of the airfoil.

For example, in certain exemplary embodiments the airfoil band includesa band flange, wherein the band flange defines a first side facing thecore air flowpath and an opposite second side at least partiallydefining the cavity, and wherein the sealing element is configured tocontact the second side of the band flange.

For example, in certain exemplary embodiments the airfoil band definesan airflow passage for providing pressurized air to the cavity of theairfoil band to pressurize the cavity of the airfoil band.

For example, in certain exemplary embodiments the sealing element is asubstantially planar member.

For example, in certain exemplary embodiments the airfoil band defines aband slot extending to the cavity defined by the airfoil band, andwherein the sealing element further extends at least partially into theband slot of the airfoil band.

In certain exemplary embodiments the first end of the airfoil includesan airfoil flange positioned at least partially within the cavity,wherein the airfoil band includes a band flange at least partiallydefining the cavity, wherein the band flange defines a first side facingthe core air flowpath and an opposite second side, and wherein thesealing element is configured to contact the second side of the bandflange and the airfoil flange.

For example, in certain exemplary embodiments the airfoil defines apressure side and a suction side, wherein the airfoil flange is aT-shaped flange including a first portion extending from the pressureside and a second portion extending from the suction side, and whereinthe sealing element is configured to form a seal between the firstportion of the airfoil flange and the band flange, between the secondportion of the airfoil flange into the band flange, or both.

In certain exemplary embodiments the airfoil band includes an innersurface at least partially within the cavity, and wherein the sealingelement is configured to contact and form a seal between the first endof the airfoil and the inner surface of the airfoil band.

For example, in certain exemplary embodiments the first end of theairfoil defines a channel for at least partially receiving the sealingelement.

In certain exemplary embodiments the cavity defines an opening in theflowpath surface, wherein the opening of the cavity defines a shapelarger than a cross-sectional shape of a portion of the first end of theairfoil extending through the opening of the cavity.

In certain exemplary embodiments the variable guide vane assembly is avariable stator vane configured for a turbine section of the gas turbineengine.

In certain exemplary embodiments the variable guide vane assemblyfurther includes an actuation member operable with the airfoil to rotateat least a portion of the airfoil.

In another exemplary embodiment of the present disclosure, a gas turbineengine is provided. The gas turbine engine includes a compressorsection; a combustion section; and a turbine section. The compressorsection, combustion section, and turbine section our in serial floworder and define at least in part a core air flowpath. The turbinesection includes a variable guide vane assembly, the variable guide vaneassembly including an airfoil band defining a flowpath surface exposedto the core air flowpath and a cavity; and an airfoil including a firstend extending at least partially into the cavity of the airfoil band andan opposite second end. The airfoil extending generally along an axisbetween the first end and the second end and being moveable generallyabout the axis relative to the airfoil band. The variable guide vaneassembly also including a sealing element operable to form a sealbetween the first end of the airfoil and the airfoil band.

In certain exemplary embodiments the sealing element is positioned atleast partially within the cavity of the airfoil band.

For example, in certain exemplary embodiments the airfoil defines aleading edge and a trailing edge, wherein the first end of the airfoildefines an airfoil slot extending in a direction generally from theleading edge to the trailing edge, and wherein the sealing elementextends at least partially into the airfoil slot of the airfoil.

For example, in certain exemplary embodiments the airfoil defines apressure side and a suction side, wherein the airfoil slot of theairfoil extends from the pressure side to the suction side, and whereinthe sealing element extends through the airfoil slot of the airfoil.

For example, in certain exemplary embodiments the airfoil band includesa band flange, wherein the band flange defines a first side facing thecore air flowpath and an opposite second side at least partiallydefining the cavity, and wherein the sealing element is configured tocontact the second side of the band flange.

These and other features, aspects and advantages of the presentinvention will become better understood with reference to the followingdescription and appended claims. The accompanying drawings, which areincorporated in and constitute a part of this specification, illustrateembodiments of the invention and, together with the description, serveto explain the principles of the invention.

BRIEF DESCRIPTION OF THE DRAWINGS

A full and enabling disclosure of the present invention, including thebest mode thereof, directed to one of ordinary skill in the art, is setforth in the specification, which makes reference to the appendedfigures, in which:

FIG. 1 is a schematic cross-sectional view of an exemplary gas turbineengine according to various embodiments of the present subject matter;

FIG. 2 is a side cross-sectional view of a compressor section, acombustion section, and a high pressure turbine section of the gasturbine engine shown in FIG. 1;

FIG. 3 is a perspective view of a second stage of variable guide vanesin a turbine section of the gas turbine engine shown in FIG. 1;

FIG. 4 is a perspective view of an airfoil of a variable guide vaneassembly including an embodiment of the present disclosure;

FIG. 5 is a cross-sectional view taken along line 5-5 of FIG. 3;

FIG. 6 is a cross-sectional view of a section of a variable guide vaneassembly in accordance with another exemplary embodiment of the presentdisclosure;

FIG. 7 is a cross-sectional view of a section of a variable guide vaneassembly in accordance with still another exemplary embodiment of thepresent disclosure; and

FIG. 8 is a cross-sectional view of a section of a variable guide vaneassembly in accordance with yet another exemplary embodiment of thepresent disclosure.

DETAILED DESCRIPTION

Reference will now be made in detail to present embodiments of theinvention, one or more examples of which are illustrated in theaccompanying drawings. The detailed description uses numerical andletter designations to refer to features in the drawings. Like orsimilar designations in the drawings and description have been used torefer to like or similar parts of the invention.

As used herein, the terms “first”, “second”, and “third” may be usedinterchangeably to distinguish one component from another and are notintended to signify location or importance of the individual components.

The terms “forward” and “aft” refer to relative positions within a gasturbine engine or vehicle, and refer to the normal operational attitudeof the gas turbine engine or vehicle. For example, with regard to a gasturbine engine, forward refers to a position closer to an engine inletand aft refers to a position closer to an engine nozzle or exhaust.

The terms “upstream” and “downstream” refer to the relative directionwith respect to fluid flow in a fluid pathway. For example, “upstream”refers to the direction from which the fluid flows, and “downstream”refers to the direction to which the fluid flows.

The terms “coupled,” “fixed,” “attached to,” and the like refer to bothdirect coupling, fixing, or attaching, as well as indirect coupling,fixing, or attaching through one or more intermediate components orfeatures, unless otherwise specified herein.

The singular forms “a”, “an”, and “the” include plural references unlessthe context clearly dictates otherwise.

Approximating language, as used herein throughout the specification andclaims, is applied to modify any quantitative representation that couldpermissibly vary without resulting in a change in the basic function towhich it is related. Accordingly, a value modified by a term or terms,such as “about”, “approximately”, and “substantially”, are not to belimited to the precise value specified. In at least some instances, theapproximating language may correspond to the precision of an instrumentfor measuring the value, or the precision of the methods or machines forconstructing or manufacturing the components and/or systems. Forexample, the approximating language may refer to being within a 10percent margin.

Here and throughout the specification and claims, range limitations arecombined and interchanged, such ranges are identified and include allthe sub-ranges contained therein unless context or language indicatesotherwise. For example, all ranges disclosed herein are inclusive of theendpoints, and the endpoints are independently combinable with eachother.

Referring now to the drawings, FIG. 1 is a schematic cross-sectionalview of a gas turbine engine 100 in accordance with an exemplaryembodiment of the present disclosure. More particularly, for theembodiment of FIG. 1, the gas turbine engine 100 is an aeronautical,high-bypass turbofan jet engine configured to be mounted to an aircraft,such as in an under-wing configuration or tail-mounted configuration. Asshown in FIG. 1, the gas turbine engine 100 defines an axial directionA1 (extending parallel to or coaxial with a longitudinal centerline 102provided for reference), a radial direction R1, and a circumferentialdirection C1 (i.e., a direction extending about the axial direction A1;see FIG. 3). In general, the gas turbine engine 100 includes a fansection 104 and a turbomachine 106 disposed downstream from the fansection 104.

The exemplary turbomachine 106 depicted generally includes asubstantially tubular outer casing 108 that defines an annular inlet110. The outer casing 108 encases, in serial flow relationship, acompressor section 112 including a first, booster or LP compressor 114and a second, HP compressor 116; a combustion section 118; a turbinesection 120 including a first, HP turbine 122 and a second, LP turbine124; and a jet exhaust nozzle section 126. A HP shaft or spool 128drivingly connects the HP turbine 122 to the HP compressor 116. ALPshaft or spool 130 drivingly connects the LP turbine 124 to the LPcompressor 114. The compressor section, combustion section 118, turbinesection, and jet exhaust nozzle section 126 together define a core airflowpath 132 through the turbomachine 106.

Referring still the embodiment of FIG. 1, the fan section 104 includes avariable pitch fan 134 having a plurality of fan blades 136 coupled to adisk 138 in a circumferentially spaced apart manner. As depicted, thefan blades 136 extend outwardly from disk 138 generally along the radialdirection R. Each fan blade 136 is rotatable relative to the disk 138about a pitch axis P by virtue of the fan blades 136 being operativelycoupled to a suitable actuation member 140 configured to collectivelyvary the pitch of the fan blades 136, e.g., in unison. The fan blades136, disk 138, and actuation member 140 are together rotatable about thelongitudinal centerline 102 by LP shaft 130 across a power gear box 142.The power gear box 142 includes a plurality of gears for stepping downthe rotational speed of the LP shaft 130 to a more efficient rotationalfan speed.

Referring still to the exemplary embodiment of FIG. 1, the disk 138 iscovered by rotatable front nacelle 144 aerodynamically contoured topromote an airflow through the plurality of fan blades 136.Additionally, the exemplary fan section 104 includes an annular fancasing or outer nacelle 146 that circumferentially surrounds the fan 134and/or at least a portion of the turbomachine 106. Moreover, for theembodiment depicted, the nacelle 146 is supported relative to theturbomachine 106 by a plurality of circumferentially spaced outlet guidevanes 148. Further, a downstream section 150 of the nacelle 146 extendsover an outer portion of the turbomachine 106 so as to define a bypassairflow passage 152 therebetween.

During operation of the gas turbine engine 100, a volume of air 154enters the gas turbine engine 100 through an associated inlet 156 of thenacelle 146 and/or fan section 104. As the volume of air 154 passesacross the fan blades 136, a first portion of the air 154 as indicatedby arrows 158 is directed or routed into the bypass airflow passage 152and a second portion of the air 154 as indicated by arrow 160 isdirected or routed into the LP compressor 114. The pressure of thesecond portion of air 160 is then increased as it is routed through thehigh pressure (HP) compressor 116 and into the combustion section 118.

Referring still to FIG. 1, the compressed second portion of air 160 fromthe compressor section mixes with fuel and is burned within thecombustion section 118 to provide combustion gases 162. The combustiongases 162 are routed from the combustion section 118 along the hot gaspath 174, through the HP turbine 122 where a portion of thermal and/orkinetic energy from the combustion gases 162 is extracted via sequentialstages of HP turbine stator vanes 164 that are coupled to the outercasing 108 and HP turbine rotor blades 166 that are coupled to the HPshaft or spool 128, thus causing the HP shaft or spool 128 to rotate,thereby supporting operation of the HP compressor 116. The combustiongases 162 are then routed through the LP turbine 124 where a secondportion of thermal and kinetic energy is extracted from the combustiongases 162 via sequential stages of LP turbine stator vanes 168 that arecoupled to the outer casing 108 and LP turbine rotor blades 170 that arecoupled to the LP shaft or spool 130, thus causing the LP shaft or spool130 to rotate, thereby supporting operation of the LP compressor 114and/or rotation of the fan 134.

The combustion gases 162 are subsequently routed through the jet exhaustnozzle section 126 of the turbomachine 106 to provide propulsive thrust.Simultaneously, the pressure of the first portion of air 158 issubstantially increased as the first portion of air 158 is routedthrough the bypass airflow passage 152 before it is exhausted from a fannozzle exhaust section 172 of the gas turbine engine 100, also providingpropulsive thrust. The HP turbine 122, the LP turbine 124, and the jetexhaust nozzle section 126 at least partially define a hot gas path 174for routing the combustion gases 162 through the turbomachine 106.

It will be appreciated that the exemplary gas turbine engine 100depicted in FIG. 1 is by way of example only, and that in otherexemplary embodiments, the gas turbine engine 100 may have any othersuitable configuration. For example, in other embodiments, the gasturbine engine 100 may be a variable bypass engine, may not include apower gearbox 142, may include a fixed-pitch fan, etc. Additionally, oralternatively, aspects of the present disclosure may be utilized withany other suitable aeronautical gas turbine engine, such as a turboshaftengine, turboprop engine, turbojet engine, etc. Further, aspects of thepresent disclosure may further be utilized with any other land-based gasturbine engine, such as a power generation gas turbine engine, or anyaeroderivative gas turbine engine, such as a nautical gas turbineengine.

FIG. 2 provides a side cross-sectional view of the compressor section112, combustion section 118, and the turbine section 120 of theturbomachine 106 of FIG. 1. More specifically, the rear end of the HPcompressor 116, the combustor section 118, and the forward end of the HPturbine 122 are illustrated.

Compressed air 176 exits the HP compressor 116 through a diffuser 178located at the rear end or outlet of the HP compressor 116 and diffusesinto the combustion section 118. The combustion section 118 ofturbomachine 106 is annularly encased by radially inner and outercombustor casings 180, 182. The radially inner combustor casing 180 andthe radially outer combustor casing 182 both extend generally along theaxial direction A1 and surround a combustor assembly 184 in annularrings. The inner and outer combustor casings 180, 182 are joinedtogether at annular diffuser 178 at the forward end of the combustionsection 118.

As shown, the combustor assembly 184 generally includes an inner liner186 extending between a rear end 188 and a forward end 190 generallyalong the axial direction A1, as well as an outer liner 192 alsoextending between a rear end 194 and a forward end 196 generally alongthe axial direction A1. The inner and outer liners 186, 192 together atleast partially define a combustion chamber 198 therebetween. The innerand outer liners 186, 192 are each attached to or formed integrally withan annular dome. More particularly, the annular dome includes an innerdome section 200 formed integrally with the forward end 190 of the innerliner 186 and an outer dome section 202 formed generally with theforward end 196 of the outer liner 192. Further, the inner and outerdome section 200, 202 may each be formed integrally (or alternativelymay be formed of a plurality of components attached in any suitablemanner) and may each extend along the circumferential direction C1 todefine an annular shape. It should be appreciated, however, that inother embodiments, the combustor assembly 184 may not include the innerand/or outer dome sections 200, 202; may include separately formed innerand/or outer dome sections 200, 202 attached to the respective innerliner 186 and outer liner 192; or may have any other suitableconfiguration.

Referring still to FIG. 2, the combustor assembly 184 further includes aplurality of fuel air mixers 204 spaced along the circumferentialdirection C1 and positioned at least partially within the annular dome.More particularly, the plurality of fuel air mixers 204 are disposed atleast partially between the outer dome section 202 and the inner domesection 200 along the radial direction R1. Compressed air 176 from thecompressor section 112 of the gas turbine engine 100 flows into orthrough the fuel air mixers 204, where the compressed air 176 is mixedwith fuel and ignited to create combustion gases 162 within thecombustion chamber 198. The inner and outer dome sections 200, 202 areconfigured to assist in providing such a flow of compressed air 176 fromthe compressor section 112 into or through the fuel air mixers 204.

As discussed above, the combustion gases 162 flow from the combustionchamber 198 into and through the turbine section 120 of the gas turbineengine 100, where a portion of thermal and/or kinetic energy from thecombustion gases 162 is extracted via sequential stages of turbinestator vanes and turbine rotor blades within the HP turbine 122 and LPturbine 124. More specifically, as is depicted in FIG. 2, combustiongases 162 from the combustion chamber 198 flow into the HP turbine 122,located immediately downstream of the combustion chamber 198, wherethermal and/or kinetic energy from the combustion gases 162 is extractedvia sequential stages of HP turbine stator vanes 164 (discussed ingreater detail below) and HP turbine rotor blades 166.

As illustrated in FIG. 2, not all compressed air 176 flows into ordirectly through the fuel air mixers 204 and into combustion chamber198. Some of the compressed air 176 is discharged into a plenum 206surrounding the combustor assembly 184. Plenum 206 is generally definedbetween the combustor casings 180, 182 and the liners 186, 192. Theouter combustor casing 182 and the outer liner 192 define an outerplenum 208 generally disposed radially outward from the combustionchamber 198. The inner combustor casing 180 and the inner liner 186define an inner plenum 210 generally disposed radially inward withrespect to the combustion chamber 198. As compressed air 176 is diffusedby diffuser 178, some of the compressed air 176 flows radially outwardinto the outer plenum 208 and some of the compressed air 176 flowsradially inward into the inner plenum 210.

The compressed air 176 flowing radially outward into the outer plenum208 flows generally axially to the turbine section 120. Specifically,the compressed air 176 flows above the HP turbine stator vanes 164 androtor blades 166. The outer plenum 208 may extend to the LP turbine 124(FIG. 1) as well.

As further shown in FIG. 2, for the embodiment depicted, the HP turbine122 includes a first stage 212 of turbine stator vanes 164 and a secondstage 214 of turbine stator vanes 164 (as well as a first and secondstage of turbine rotor blades 166). Moreover, for the embodimentdepicted, the second stage 214 of turbine stator vanes 164 is of avariable configuration, such that the second stage 214 of turbine statorvanes 164 includes a plurality of variable guide vane assemblies 216.

Referring now also to FIG. 3, providing a perspective view of aplurality of the exemplary variable guide vane assemblies 216 of thesecond stage 214 of turbine stator vanes 164, the plurality of variableguide vane assemblies 216 are spaced generally along the circumferentialdirection C1 of the gas turbine engine 100. Additionally, each of thevariable guide vane assemblies 216 includes an airfoil 218 extendinggenerally along an axis 220 between a first, outer end 222 (i.e., outerend along the radial direction R1) and an opposite second, inner end 224(i.e., inner end along the radial direction R1). For the embodimentdepicted, the axis 220 of each airfoil 218 is generally aligned with theradial direction R1 of the gas turbine engine 100. Moreover, eachvariable guide vane assembly 216 includes an outer airfoil band 226along the radial direction R1 and an inner airfoil band 228 along theradial direction R1. Accordingly, it will be appreciated that the outerend 222 of the airfoil 218 is positioned adjacent to the radially outerairfoil band 226, and the inner end 224 of the airfoil 218 is positionedadjacent to the radially inner airfoil band 228. Additionally, the innerairfoil band 228 defines a flowpath surface 230 and the outer airfoilband 226 also defines a flowpath surface 230 (see FIG. 2)—the flowpathsurface 230 of the inner airfoil band 228 and the flowpath surface 230of the outer airfoil band 226 each at least partially defining the coreair flowpath 132 through the gas turbine engine 100.

As is further depicted, each variable guide vane assembly 216 includesan actuation member 232 operable with the respective airfoil 218 forrotating at least a portion of the respective airfoil 218 along itsrespective axis 220. For the embodiment depicted, each of the actuationmembers 232 the plurality of variable guide vane assemblies 216 areconfigured together as a single actuation member assembly 234, such thatthe actuation member assembly 234 may actuate the plurality of airfoils218 of the plurality of variable guide vane assemblies 216, e.g., inunison. More particularly, for the embodiment depicted, each actuationmember 232 generally includes a hub 235 and an arm 236, and theactuation member assembly 234 includes a ring 238 connected to the arms236 of each of the actuation members 232 to move the hubs 235 of each ofthe actuation members 232 together, e.g., in unison. The hubs 235 of theactuation members 232 may be coupled a respective airfoil 218 forrotating the respective airfoil 218. However, in other embodiments, theindividual actuation members 232 may move at least partiallyindependently from the other actuation members 232.

Furthermore, for the embodiment depicted, the inner airfoil bands 228 ofadjacent variable guide vane assemblies 216 are coupled together to forma substantially continuous inner airfoil band assembly, and similarly,the outer airfoil bands 226 of adjacent variable guide vane assemblies216 are coupled together to form a substantially continuous outerairfoil band assembly. However, in other exemplary embodiments, theinner and outer airfoil bands 226, 228 of the plurality of variableguide vane assemblies 216 may be configured in any other suitablemanner. Further, in other exemplary embodiments, the actuation memberassembly 234 may be configured in any other suitable manner foractuating the plurality of airfoils 218. Further, still, in otherexemplary embodiments, the airfoils 218 of the variable guide vaneassemblies 216 may be configured in any other suitable manner forvarying an effective flow angle across the respective airfoils 218. Forexample, in other exemplary embodiments, the entire airfoil 218 of eachvariable guide vane assembly 216 may not be movable, and instead, eachairfoil 218 may include a tail section (e.g., at a trailing edge 246)configured to rotate about an axis 220 to vary an effective flow angleacross the respective airfoil 218.

Referring now to FIGS. 4 and 5, it will be appreciated that at least oneof the outer end 222 of the airfoil 218 or the inner end 224 of theairfoil 218 of each variable guide vane assembly 216 is configured withthe respective outer airfoil band 226 and inner airfoil band 228 toprevent airflow from an upstream location or high pressure location ofthe variable guide vane assemblies 216 from “leaking” to a downstreamlocation or low pressure location. FIG. 4 provides a perspective view ofan outer end 222 of an airfoil 218 of an exemplary variable guide vaneassembly 216 and FIG. 5 provides a cross-sectional view of the outer end222 of the exemplary airfoil 218 of FIG. 4 configured with an outerairfoil band 226 of the variable guide vane assembly 216. The variableguide vane assembly 216 of FIGS. 4 and 5 may be the same variable guidevane assembly 216 described above with reference to FIGS. 1 through 3.As is depicted, the variable guide vane assembly 216 of FIGS. 4 and 5further includes a sealing element 240 operable to form a seal betweenthe outer end 222 of the airfoil 218 and the outer airfoil band 226.Notably, although not depicted, the variable guide vane assembly 216 mayfurther include another sealing element operable to form a seal betweenthe inner end 224 of the airfoil 218 the inner airfoil band 228.

As is depicted, the outer end 222 of the airfoil 218 defines an airfoilslot 242. More particularly, the airfoil 218 generally includes aleading edge 244 and an opposite trailing edge 246, as well as apressure side 248 and an opposite suction side 250. The airfoil slot 242defined by the outer end 222 of the airfoil 218 extends in a directiongenerally from the leading edge 244 to the trailing edge 246, andfurther extends from the pressure side 248 to the suction side 250. Moreparticularly, the airfoil slot 242 extends through the outer end 222 ofthe airfoil 218 from the pressure side 248 to the suction side 250, andfurther extends substantially from the leading edge 244 to the trailingedge 246. However, in other exemplary embodiments, the airfoil slot 242may not extend completely from the pressure side 248 to the suction side250 (e.g., see FIG. 3), and may not extend substantially from theleading edge 244 to a trailing edge 246. For example, the airfoil 218generally defines a camber line 252. The airfoil slot 242 may extendalong at least fifty percent of the camber line 252, such as at leastabout sixty percent of the camber line 252 such as at least aboutseventy-five percent of camber line 252.

Additionally, as is depicted, the sealing element 240 extends at leastpartially into the airfoil slot 242 of the airfoil 218. Moreparticularly, for the embodiment depicted the sealing element 240extends completely through the airfoil slot 242 of the airfoil 218(i.e., from the pressure side 248 through the suction side 250).Additionally, the sealing element 240 is a substantially planar member.Notably, the airfoil slot 242 defines a height 254 generally along theaxis 220 of the airfoil 218 and the sealing element 240 defines athickness 256. The thickness 256 of the sealing element 240 is less thanthe height 254 of the airfoil slot 242, allowing the sealing element 240to move relative to the airfoil 218 generally along the axis 220 of theairfoil 218. As will be discussed in greater detail below, such mayallow for the sealing element 240 to form a better seal with the outerairfoil band 226. For example, in certain embodiments, the thickness 256of the sealing element 240 may be at least about ten percent less thanthe height 254 of the airfoil slot 242, such as at least about twentypercent less than the height 254 of the airfoil slot 242, such as atleast about thirty percent less than the height 254 of the airfoil slot242, such as at least about fifty percent less than the height 254 ofthe airfoil slot 242. As used herein, the terminology “variable B beingX percent less than variable A” refers to the variable B being equal tovariable A minus variable A times X.

Referring now particularly to FIG. 5, which is a view along Line 5-5 ofFIG. 4, it will be appreciated that the outer airfoil band 226 furtherdefines a cavity 258 having an opening 260 in the flowpath surface 230of the outer airfoil band 226. The outer end 222 of the airfoil 218extends at least partially into the cavity 258 of the outer airfoil band226, and similarly, the sealing element 240 is positioned at leastpartially within the cavity 258 of the outer airfoil band 226. Moreparticularly, for the embodiment depicted, the sealing element 240 ispositioned completely within the cavity 258 of the outer airfoil band226. Moreover, it will be appreciated that for the embodiment depicted,the opening 260 of the cavity 258 defines a shape larger than across-sectional shape of a portion of the outer end 222 of the airfoil218 extending through the opening 260 of the cavity 258. Such may allowfor the airfoil 218 to rotate about the axis 220 relative to the outerairfoil band 226 when actuated by the actuation member 232.

Furthermore, as is also depicted in FIG. 5, for the embodiment depicted,the outer airfoil band 226 further includes a band flange 262, with theband flange 262 defining a first side 264 facing the core air flowpath132 and an opposite second side 266 at least partially defining thecavity 258. The band flange 262 extends at least partially around theopening 260 of the cavity 258. Additionally, the first side 264 of theband flange 262 is substantially continuous with the flowpath surface230 of the outer airfoil band 226. In order to prevent airflow fromwithin the core air flowpath 132 from leaking from an upstream location,or pressure side 248 location, of the airfoil 218 to a downstreamlocation, or suction side 250 location, of the airfoil 218, the sealingelement 240 is configured to contact the second side 266 of the bandflange 262, creating a seal therewith. Moreover, for the embodimentdepicted, the sealing element 240 is configured to contact an innersurface 268 of the airfoil slot 242 to further form a seal therewith.

Notably, in order to increase the effectiveness of the sealing element240, the variable guide vane assembly 216 is configured to pressurizethe cavity 258 to further press the sealing element 240 against thesecond side 266 of the band flange 262 and the inner surface 268 of theairfoil slot 242. More particularly, for the exemplary embodimentdepicted, the outer airfoil band 226 defines an airflow passage 270 forproviding pressurized air to the cavity 258 of the outer airfoil band226 to pressurize the cavity 258 of the outer airfoil band 226. Moreparticularly, for the exemplary embodiment depicted, the outer airfoilband 226 defines a plurality of airflow passages 270 for providingpressurized air to the cavity 258 of the outer airfoil band 226 topressurize the cavity 258 of the outer airfoil band 226. In certainexemplary embodiments, the pressurization air may be compressed air 176received from the outer plenum 208. However, in other exemplaryembodiments, the pressurized air may be taken from any other suitablehigh pressure air source.

It will be appreciated that although the discussion herein is directedprimarily to the outer end 222 of the airfoil 218 being positionedwithin the cavity 258 of the outer airfoil band 226 with the sealingelement 240 being operable to form a seal between the outer end 222 ofthe airfoil 218 in the outer airfoil band 226, in other exemplaryembodiments the inner end 224 of the airfoil 218 and inner airfoil band228 may be configured in a similar manner. For example, in otherexemplary embodiments, the inner airfoil band 228 may also define acavity having an opening in the flowpath surface 230 with the inner end224 of the airfoil 218 positioned at least partially within the cavityof the inner airfoil band 228. Further, a sealing element may beprovided and positioned at least partially within the cavity of theinner airfoil band 228 and operable to form a seal between the inner end224 of the airfoil 218 and the inner airfoil band 228.

Inclusion of a variable guide vane assembly in accordance with one ormore of the exemplary embodiments provided herein may results in a moreefficient gas turbine engine as such a variable guide vane assembly mayreduce an amount of air leakage of the airflow through the core airflowpath from a high pressure area to a low pressure area around anouter end and/or an inner end of an airfoil of the variable guide vaneassembly.

It should be appreciated, however, that in other exemplary embodiments,the variable guide vane assembly 216 may be configured in any othersuitable manner for reducing a leakage of the airflow through the coreair flowpath 132 around the airfoil 218.

For example, referring now to FIG. 6, a cross-sectional view of an outerend 222 of an airfoil 218 positioned within an outer airfoil band 226 ofa variable guide vane assembly 216 in accordance with another exemplaryembodiment of the present disclosure is provided. The exemplary variableguide vane assembly 216 of FIG. 6 may be configured in substantially thesame manner as the exemplary variable guide vane assembly 216 describedabove with reference to FIGS. 2 through 5, and the view of FIG. 6 isfrom the same perspective as the view from FIG. 5. For example, thevariable guide vane assembly 216 may generally include an outer airfoilband 226 defining a flowpath surface 230 and a cavity 258, with thecavity 258 including an opening 260 in the flowpath surface 230. Thevariable guide vane assembly 216 may further include an airfoil 218having an outer end 222 extending at least partially into the cavity 258of the outer airfoil band 226 and a sealing element 240 operable to forma seal between the outer end 222 of the airfoil 218 and the outerairfoil band 226. Moreover, the outer end 222 of the airfoil 218 definesa leading edge 244 (see, e.g., embodiment of FIG. 4), a trailing edge246 (see, e.g., embodiment of FIG. 4), a pressure side 248, a suctionside 250, and an airfoil slot 242 extending in a direction generallyfrom the leading edge 244 to the trailing edge 246 and from the pressureside 248 to the suction side 250. The sealing element 240 extends atleast partially into the airfoil slot 242 of the airfoil 218.

However, for the exemplary aspect of FIG. 6, the outer airfoil band 226further defines a band slot 272 extending to the cavity 258 defined bythe outer airfoil band 226. The sealing element 240 is configured as aplanar member extending at least partially into the band slot 272 of theouter airfoil band 226 and further at least partially into the airfoilslot 242. More particularly, for the embodiment depicted, the band slot272 extends around the outer end 222 of the airfoil 218, such that theband slot 272 is positioned on both the pressure side 248 and thesuction side 250 of the airfoil 218, and such that the sealing element240 forms a seal around the outer end 222 of the airfoil 218 on both thepressure side 248 and the suction side 250.

Moreover, it will be appreciated that for the embodiment depicted, theairfoil slot 242 does not extend completely from the pressure side 248to the suction side 250 of the airfoil 218 for the exemplary embodimentdepicted. Accordingly, the airfoil slot 242 depicted includes a pressureside portion and a suction side portion, and similarly the sealingelement 240 includes a first portion positioned at least partially inthe pressure side portion of the airfoil slot 242 and a second portionpositioned at least partially in the suction side portion of the airfoilslot 242. Further, for the embodiment depicted, a height of the airfoilslot is substantially equal to a thickness of the sealing element 240.

Further, in still other exemplary embodiments, other configurations maybe provided. For example, referring now to FIGS. 7 and 8, two additionalembodiments are provided. The exemplary embodiments of FIGS. 7 and 8 mayeach also be configured in substantially the same manner as exemplaryembodiment described above with reference to FIGS. 2 through 5, and theviews of FIGS. 7 and 8 are each from the same perspective as the viewfrom FIG. 5. For example, the variable guide vane assemblies 216 ofFIGS. 7 and 8 each generally include an outer airfoil band 226 defininga flowpath surface 230 and a cavity 258, with the cavity 258 includingan opening 260 in the flowpath surface 230. The variable guide vaneassemblies 216 each further include an airfoil 218 having an outer end222 extending at least partially into the cavity 258 of the respectiveouter airfoil band 226 and a sealing element 240 operable to form a sealbetween the respective outer ends 222 of the airfoils 218 and the outerairfoil bands 226.

Referring first to exemplary embodiment of FIG. 7, for the embodimentdepicted, the outer airfoil band 226 further comprises a band flange262, with the band flange 262 defining a first side 264 facing the coreair flowpath 132 and an opposite second side 266 at least partiallydefining the cavity 258. The band flange 262 extends at least partiallyaround the opening 260 of the cavity 258. Additionally, the first side264 of the band flange 262 is substantially continuous with the flowpathsurface 230 of the outer airfoil band 226. Further, the outer end 222 ofthe airfoil 218 includes an airfoil flange 275 positioned at leastpartially within the cavity 258. For the embodiment depicted, thesealing element 240 is configured to contact the second side 266 of theband flange 262 and the airfoil flange 275 to provide a sealtherebetween. More particularly, for the embodiment depicted, theairfoil 218 defines a pressure side 248 and a suction side 250 and theairfoil flange 275 is configured as a T-shaped flange. The T-shapedflange includes a first portion 274 extending from the pressure side 248of the airfoil 218 and a second portion 276 extending from the suctionside 250 of the airfoil 218. The sealing element 240 similarly includesa first portion 278 configured to form a seal between the first portion274 of the airfoil flange 275 and the second side 266 of the band flange262, as well as a second portion 280 configured to form a seal betweenthe second portion 276 of the airfoil flange 275 and the second side 266of the band flange 262. However, in other exemplary embodiments, theairfoil flange 275 may be positioned on the other side of the bandflange 262, such that the sealing elements 278, 280 instead contact, andform a seal between, the first side 264 of the band flange 262 and theairfoil flange 275. Further, although for the embodiment depicted thesealing element 240 is configured to form a seal between the firstportion 274 of the airfoil flange 275 and the band flange 262, as wellas between the second portion 276 of the airfoil flange 275 and the bandflange 262, in other exemplary embodiments, the sealing element 240 mayinstead be configured to form a seal between only one of the firstportion 274 of the airfoil flange 275 and the band flange 262, orbetween the second portion 276 of the airfoil flange 275 and the bandflange 262

Notably, for the embodiment depicted, the first portion 278 of thesealing element 240 defines a different cross-sectional shape than thesecond portion 280 of the sealing element 240. Such may allow for thefirst and second portions 278, 280 of the sealing element 240 to providedifferent sealing thresholds (i.e., be capable of withstanding differentdifferential pressures). However, in other exemplary embodiments, thefirst and second portions 278, 280 of the sealing element 240 may havesame cross-sectional shapes.

Referring now to the exemplary embodiment of FIG. 8, for the embodimentdepicted, the outer airfoil band 226 further includes an inner surface282 at least partially within the cavity 258, and at least partiallydefining the cavity 258. Further, the outer end 222 of the airfoil 218includes a radially outer surface 284. The sealing element 240 isconfigured to contact and form a seal between the radially outer surface284 of the outer end 222 of the airfoil 218 and the inner surface 282 ofthe outer airfoil band 226. More particularly, for the embodimentdepicted, the outer end 222 of the airfoil 218, or rather, the radiallyouter surface 284 of the outer end 222 of the airflow, defines a channel286 for at least partially receiving the sealing element 240. Such achannel 286 may ensure the sealing element 240 remains in place during arotation or actuation of the airfoil 218. Notably, however, in otherexemplary embodiments, the inner surface 282 of the outer airfoil band226 may instead define the channel.

It will be appreciated, that the various sealing elements 240 describedherein may be formed of any suitable material for forming a seal betweenthe two components. Notably, for each of the exemplary embodimentdepicted, the variable guide vane assemblies 216 are configured asvariable stator vanes configured for a turbine section of the gasturbine engine (depicted as being in the HP turbine 122, but mayalternatively be in the LP turbine 124 or an intermediate turbine, ifprovided). Accordingly, the various sealing elements 240 describedherein may be capable of withstanding the relatively high temperatureswithin a respective portion of the turbine section.

Notably, however, in other exemplary embodiments, the variable guidevane assembly may instead be positioned at any other suitable locationwithin a gas turbine engine (such as, for example, a compressorsection). Additionally, although described herein as being utilized withthe gas turbine engine, in other exemplary embodiments, the variableguide vane assembly may instead be positioned within any other suitablemachine, such as within any other suitable machine including a turbineor a compressor.

This written description uses examples to disclose the invention,including the best mode, and also to enable any person skilled in theart to practice the invention, including making and using any devices orsystems and performing any incorporated methods. The patentable scope ofthe invention is defined by the claims, and may include other examplesthat occur to those skilled in the art. Such other examples are intendedto be within the scope of the claims if they include structural elementsthat do not differ from the literal language of the claims, or if theyinclude equivalent structural elements with insubstantial differencesfrom the literal languages of the claims.

What is claimed is:
 1. A variable guide vane assembly for a machinedefining a core air flowpath, the variable guide vane assemblycomprising: an airfoil band defining a surface exposed to the core airflowpath; an airfoil comprising a first end and an opposite second end,the airfoil extending along an axis between the first end and the secondend and being moveable about the axis relative to the airfoil band; anda sealing element positioned at the first end to contact and form a sealbetween the first end of the airfoil and the surface of the airfoilband.
 2. The variable guide vane assembly of claim 1, wherein the firstend of the airfoil includes an outer surface along the radial direction,and wherein the sealing element is configured to contact and form a sealbetween the radially outer surface of the outer end of the airfoil andthe surface of the airfoil band.
 3. The variable guide vane assembly ofclaim 1, wherein the first end of the airfoil defines a channel for atleast partially receiving the sealing element.
 4. The variable guidevane assembly of claim 1, wherein the machine is a gas turbine engine.5. The variable guide vane assembly of claim 4, wherein the airfoil is avariable stator vane configured for a turbine section of the gas turbineengine.
 6. The variable guide vane assembly of claim 4, wherein theairfoil is a variable stator vane configured for a high pressure turbineof a turbine section of the gas turbine engine.
 7. The variable guidevane assembly of claim 1, wherein the airfoil extends between a leadingedge and a trailing edge, and wherein the sealing element extend along alength of the airfoil between the leading edge and the trailing edge. 8.The variable guide vane assembly of claim 1, wherein the sealing elementextends from a location forward of the axis to a location aft of theaxis.
 9. The variable guide vane assembly of claim 1, wherein theairfoil band is an outer airfoil band.
 10. The variable guide vaneassembly of claim 9, wherein the sealing element is location inward ofthe surface of the airfoil band along a radial direction.
 11. A gasturbine engine comprising: a compressor section; a combustion section;and a turbine section, the compressor section, combustion section, andturbine section in serial flow order and defining at least in part acore air flowpath, the turbine section comprising a variable guide vaneassembly comprising an airfoil band defining a surface exposed to thecore air flowpath; an airfoil comprising a first end and an oppositesecond end, the airfoil extending along an axis between the first endand the second end and being moveable about the axis relative to theairfoil band; and a sealing element positioned at the first end tocontact and form a seal between the first end of the airfoil and thesurface of the airfoil band.
 12. The gas turbine engine of claim 11,wherein the first end of the airfoil includes an outer surface along theradial direction, and wherein the sealing element is configured tocontact and form a seal between the radially outer surface of the outerend of the airfoil and the surface of the airfoil band.
 13. The gasturbine engine of claim 11, wherein the first end of the airfoil definesa channel for at least partially receiving the sealing element.
 14. Thegas turbine engine of claim 11, wherein the airfoil is a variable statorvane configured for a turbine section of the gas turbine engine.
 15. Thegas turbine engine of claim 11, wherein the airfoil is a variable statorvane configured for a high pressure turbine of a turbine section of thegas turbine engine.
 16. The gas turbine engine of claim 11, wherein theairfoil extends between a leading edge and a trailing edge, and whereinthe sealing element extend along a length of the airfoil between theleading edge and the trailing edge.
 17. The gas turbine engine of claim11, wherein the sealing element extends from a location forward of theaxis to a location aft of the axis.
 18. The gas turbine engine of claim11, wherein the airfoil band is an outer airfoil band.
 19. The gasturbine engine of claim 18, wherein the sealing element is locationinward of the surface of the airfoil band along a radial direction.